1. Field of the Invention
This invention relates to rocket motors and particularly to a system for reducing, neutralizing and/or reversing the thrust vector resultant of an operating solid propellant rocket motor.
2. Description of the Prior Art
Solid propellant rocket motors consist, in principal parts, of a combustion chamber within which the propellant is stored prior to ignition, and one or more nozzles through which hot gases exhaust during combustion of the propellant. The rate of gas generation during motor operation is determined by the chemical characteristics of the propellant, the surface area history of the propellant charge as it is consumed, and the pressure within the combustion chamber. The pressure is strongly affected by the flow characteristics of the exhaust gas stream. Normally, in order to maximize the thrust generated by the stream, the total flow area at the nozzle throat is proportioned to assure supersonic flow of the stream farther out.
The thrust generated by the hot gas stream exhausting from the rocket motor, used to propel a satellite, a ballistic missile, or some other payload, normally continues until substantially all of the propellant is consumed and combustion stops. Termination and reversal of the thrust vector resultant prior to this time in the combustion history has been extremely difficult.
For satellite applications, when launch point and desired altitude are determined in advance, this difficulty is not of major consequence except in lower stages of a multistage system, where stage separation must be accomplished. If the lower stage continues to be propulsive prior to ignition of the next stage, collision can occur. Similarly in weapon system applications, the difficulty is important when the nature of the mission requires separation of the rocket motor from the payload prior to its encounter with the target, particularly when the weapon must be useful for targets near and far within the range of the system. For such weapon systems, thrust management prior to completion of burn of the propellant charge is an essential feature.
Thrust termination has been accomplished in the prior art by greatly enlarging the exhaust area available for outflow of gases generated in the combustion chamber. To this end, mechanical means for release of the rocket motor nozzle, and explosive features to cut holes in the combustion chamber, have been devised. Their effect is to diminish the combustion chamber Pressure below the level necessary for sustained combustion.
Sometimes termination alone is not enough. When separation and termination are accomplished where atmospheric drag slows the payload after separation from the rocket motor and the payload shields the motor from some of the drag, collision can still occur. Preventative measures have included separate thrusters, small rocket motors mounted on the forward end of the main motor to create an aftward thrust after the end of main motor burn; there have also been devices to create additional exhaust openings at the forward end of the main motor before or after completion of its burn. For both of these approaches, an opportunity is created that hot gases exhausted in a forward direction will damage the payload.
Proposals of techniques for thrust reversal and/or neutralization using aft end features of the rocket motor have been nearly as rare as their application. One such proposal is disclosed in U.S. Pat. No. 3,177,655 granted on Apr. 13, 1965 to Roger F. White wherein circumferentially arranged ports in the combustion chamber wall of a solid propellant rocket motor are closed by a slide or cover having a forwardly curved deflector portion, which cover normally is retained in the port closing position by a segmented metal band clamp but is releasable by the activation of explosive bolts. When released, the cover slides aft until it engages a stop on the exit cone outer surface thereby uncovering the ports. This results in a diversion of some of the combustion gases outwardly through the ports. The outward flow of combustion gases through the ports is directed by the cover deflector portion forwardly of the rocket motor and provides brief thrust in the reverse direction in opposition to the primary motor thrust until the combustion chamber pressure is reduced sufficiently to cause termination of combustion of the propellant.
U.S. Pat. No. 3,196,610 granted on July 27, 1965 to Frank P. Anderson discloses a rocket motor nozzle arrangement for producing a thrust in opposition to the motor primary forward thrust comprising auxiliary forwardly curved nozzles that are connected to the combustion chamber, each nozzle being normally closed by a rupturable diaphragm. Rupturing of the diaphragms initiates the opposing thrust.
It is desirable in the operation of some solid propellant rocket motors to have the ability to rapidly terminate the forward thrust, and additionally, to produce a net reverse or negative thrust. Failure to terminate the forward velocity rapidly enough may cause a missile to miss the target. For example, upon separation of the rocket motor from its missile payload, which may be a warhead, it is desirable not only to terminate the forward thrust but also to produce a net reverse or negative thrust, that is, to reverse the thrust vector resultant.
It is further desirable in the operation of a solid propellant rocket motor to have the ability to predetermine the amount of reverse thrust and the time required to reach a reverse or negative thrust vector resultant thereby to provide a reduction, neutralization, or reversal of the thrust vector resultant, as desired.